1. Field of the Invention
The present invention relates generally to gas turbine engine blades and, more particularly, to turbine blade cooling and turbine blade platforms.
2. Discussion of the Background Art
A gas turbine engine includes a compressor for pressurizing air which is channeled to a combustor wherein it is mixed with fuel and ignited for generating hot combustion gas. The combustion gas flows downstream through one or more turbine stages which extract energy therefrom for producing work. A typical turbine blade includes a dovetail disposed in a complementary dovetail slot in a perimeter of a disk of a turbine rotor for securing the blade thereto. A shank extends radially outwardly from the dovetail to a platform which defines a radially inner flowpath for the combustion gas. The airfoil extends radially outwardly from the platform for extracting energy from the combustion gas for rotating the disk and producing power.
Turbine blades are directly exposed to the hot combustion gases and are typically cooled using a portion of compressed air bled from the compressor and channeled through a cooling circuit within the airfoil of the blade. For high performance gas turbines having substantially high combustion gas temperature, the turbine blade utilizes various film cooling holes over an airfoil thereof for providing thin films of cooling air to protect the airfoil from the hot combustion gas which flows thereover.
The blade may be cooled by variously configured cooling circuits and cooling holes through the airfoil. The cooling circuit extends from the bottom of the dovetail which first receives the coolant channeled thereto, and extends upwardly through the dovetail, shank, platform, and airfoil. The cooling circuit itself provides effective cooling of the dovetail, shank, and platform since they are disposed radially inwardly of the combustion gas flowpath.
The hottest combustion gas typically flows near the mid-span region of the airfoil and first engages the airfoil along its leading edge and pressure and suction sides. Accordingly, the leading edge and pressure and suction sides of the airfoil are typically provided with suitable film cooling holes for maximizing the cooling thereof for effecting a suitably long useful life of the blade during operation.
The efficiency of the gas turbine engine may be further increased by increasing the temperature of the combustion gas, which correspondingly increases the difficulty of cooling the turbine blade. Undesirable exhaust emissions may be reduced by providing substantially flat temperature profiles for the combustion gas exiting the combustor which reduces the center-peaked temperature and effects a more radially uniform, yet high temperature, profile. This further increases the complexity of adequately cooling the turbine blade since the heat load is being distributed more uniformly from the root to tip of the airfoil.
In particular, conventional blade platforms are relatively thin plate members which have no internal cooling circuits therein. The platform is conventionally cooled solely by the coolant channeled upwardly through the shank and center of the platform into the airfoil. Accordingly, conventional uncooled blade platforms are subject to substantial thermal distress in advanced, low emission turbine engines. However, since the platforms are relatively thin and project outwardly from the airfoil, providing cooling circuits therein, while maintaining suitable strength thereof is a significant problem.
New high performance gas turbines are being designed with lower solidity or less airfoils than have been used in the past. These turbine blades require more airflow turning for each airfoil from the leading edge to the trailing edge. The larger turning results in a longer or wider platform overhang as measured from the shank. This, in turn, requires an increase in the thickness of the platform in order to accommodate or withstand the centrifugal force loading of the platform under high rotating speeds of the rotor. The platforms are subject to heating from the main gas flowpath above the platform and cooling by the rotor cooling air under the platform. The increased platform thickness will increase the undesirable weight and platform temperature. It is, therefore, desirable to have a design which can avoid or reduce the increase of the platform thickness and yet still can maintain the mechanical strength under the high rotational speed condition. It is also desirable to have a platform design that does not require cooling holes or passages therethrough.